Refractory ceramic component for a gas turbine engine

ABSTRACT

A refractory ceramic component for a gas turbine engine is employed that is more cost effective than typical components used in the gas turbine engine. The refractory ceramic component may be a refractory ceramic liner that is easily replaceable. The refractory ceramic liner may be a unitary construction or made of numerous bricks that are interlocked. The ceramic used is a refractory oxide material.

BACKGROUND 1. Field

Disclosed embodiments are generally related to components in gas turbineengines.

2. Description of the Related Art

Gas turbines comprise a casing or cylinder for housing a compressorsection, a combustion section, and a turbine section. A supply of air iscompressed in the compressor section and directed into the combustionsection. The compressed air enters the combustion inlet and is mixedwith fuel. The air/fuel mixture is then combusted to produce hightemperature and high pressure (working) gas. This working gas thentravels through the transition and into the turbine section of theturbine.

The turbine section may comprise rows of vanes which direct the workinggas to the airfoil portions of the turbine blades. The working gastravels through the turbine section, causing the turbine blades torotate, thereby turning a rotor associated therewith.

Higher efficiency of a combustion turbine can be achieved by increasingthe temperature of the working gas flowing through the combustionsection to as high a temperature as is practical. The aggressive hotgas, however, can degrade various metal turbine components, such as thecombustor, transition ducts, vanes, ring segments, and turbine blades asit flows through the turbine.

For this reason, strategies have been developed to protect turbinecomponents from extreme temperatures, such as the development andselection of high temperature materials adapted to withstand theseextreme temperatures and cooling strategies to keep the componentsadequately cooled during operation. Superalloys with additionalprotective coatings are commonly used for hot gas path components of gasturbines. In view of the substantial and longstanding development in thearea of superalloys, further increases in the temperature capability ofsuperalloys has become more difficult.

Ceramic matrix composite (CMC) materials have been developed andincreasingly utilized in gas turbine engines. Typically, CMC materialsinclude a ceramic or a ceramic matrix material, either of which hosts aplurality of reinforcing fibers. The fibers may have a predeterminedorientation to provide the CMC materials with additional mechanicalstrength. Generally, (fiber reinforced) ceramic matrix composites aremanufactured by the infiltration of a matrix slurry (e.g., alumina,mullite, silicon-containing polymers, molten silicon, or the like) intoa fiber preform. While these materials may offer a higher temperatureresistance than superalloys, fiber grains of the CMC may coarsen andresult in reduced strength over time. In addition, matrix graincoarsening can result in CMC embrittlement leading to a propensity forcracking and crack propagation as firing temperatures increase.

While both of the above strategies are frequently employed, their usagecan necessitate additional cooling strategies and can be undesirablyexpensive for certain applications.

SUMMARY

Briefly described, aspects of the present disclosure relate to acomponent for a gas turbine engine.

An aspect of present disclosure may be a gas turbine engine comprising acombustor basket. A cone may be connected to the combustor basket. Thecone comprises a liner; and a shell surrounding the liner, wherein theliner is formed from a refractory oxide ceramic material and the shellis formed from metal.

Another aspect of the present disclosure may be a component for a gasturbine engine. The gas turbine engine may comprise a liner; and a shellsurrounding the inner layer, wherein the inner layer is formed from arefractory oxide ceramic material and the shell is formed from metal.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a side view of a gas turbine engine employing a cone usingrefractory ceramic bricks to form the liner.

FIG. 2 is a cut away view of a cone formed with the refractory ceramicbricks.

FIG. 3 is another view of the refractory ceramic bricks illustrating themating of the ceramic bricks.

FIG. 4 is a view of a refractory ceramic brick and an interlayer.

FIG. 5 is a view of the cone and its connection to the combustor.

FIG. 6 is cross sectional view of a cone that uses a metal springlocated between the shell and the liner.

DETAILED DESCRIPTION

To facilitate an understanding of embodiments, principles, and featuresof the present disclosure, they are explained hereinafter with referenceto implementation in illustrative embodiments. Embodiments of thepresent disclosure, however, are not limited to use in the describedsystems or methods.

A ceramic liner is proposed for use in the gas turbine engines that isable to provide the protection desired for the gas turbine engine. Thisis a ceramic liner that avoids the costs associated with CMC. FIG. 1shows a portion of gas turbine engine 100 having a cone 10 that utilizesthe proposed ceramic liner 12. The cone 10 is connected to an integratedexit piece (IEP) 14 and transmits the working fluids through the gasturbine engine 100. While the application discusses the proposed ceramicliner 12 with reference to a cone 10, it should be understood that theceramic liner 12 may be used for other gas turbine engine components,such as transitions, blade tip seals, and interstage turbine ducts.

FIG. 1 shows the cone 10 connected to the combustor basket 8. The cone10 is formed by the liner 12 and the shell 13. Combustor 6 producesworking gases within the combustor basket 8. The working gases then flowinto the cone 10. As discussed above these working gases are very hot.The liner 12 is formed from a refractory ceramic material. Shell 13surrounds the liner 12. The shell 13 is made out of metal. An air gap 9may be present between the liner 12 and the shell 13. If an air gap 9 isutilized, air may flow through the air gap 9, which can advantageouslyassist in cooling the liner 12 and the shell 13. Alternately, the airgap can be maintained as a stagnant gap and utilized to promote radiantheat transfer between the liner 12 and shell 13.

FIG. 2 is a cut away view of the cone 10. The cone 10 is formed withrefractory ceramic bricks 15. While a plurality of bricks 15 are shownand discussed throughout the application, it is also possible to formthe liner 12 as a unitary piece. The bricks 15 shown in FIG. 2 areassembled and each of the bricks 15 formed to obtain the desired shapeof the cone 10. The desired shape of the cone 10 may change depending onthe type of gas turbine engine in which the cone 10 is being used.

Forming the liner 12 with a plurality of bricks 15 makes it easier toservice and/or replace the liner 12. For example, in the event that aportion of the liner 12, such as one of the bricks 15, has undergonesome type of damage or has simply outlived its natural life span thatparticular brick 15 may be serviced. The liner 12 may provide furthersavings in life cycle cost with only the liner 12 needing to be replacedat combustion/transition intervals, instead of the entire cone 10.

Another advantage of using a plurality of bricks 15 to form the liner 12relates to costs of using the bricks 15. The costs of the bricks 15 madefrom refractory oxide ceramic material are relatively inexpensivecompared with other ceramic materials, or forming the liner 12 out ofCMC.

Additional savings can be achieved by using bricks 15 formed fromrefractory oxide ceramic material. The refractory oxide ceramic materialis able to handle higher temperatures without cooling. When used in thegas turbine engine 100 savings can be achieved by not having to provideadditional cooling features that would be needed to cool metalcomponents and part.

The refractory ceramic bricks 15 form a liner 12 that is heavier thantypically used liners. Each of the bricks 15 has an inner surface 17 andan outer surface 19. The distance d1 (i.e. the thickness) between apoint on the inner surface 17 and a point on the outer surface 19 may begreater than 20 mm, and preferably greater than 25 mm. A range for thedistance d1 may be between 20-30 mm. The thicker or greater the distanced1, the more heat protection liner 12 will provide. The thickness ordistance d1 also compensates for the lower durability that the bricks 15made of oxide ceramic may have. The overall thickness of the refractoryceramic bricks 15 also distinguishes the liner 12 from other types ofliners.

As indicated above, the use of bricks 15 reduces the need for coolingair since the bricks 15 can withstand higher operating temperatures,such as those greater than 1400° C. This greater than the temperaturesthat other types of materials can typically withstand. By usingmaterials that permit the gas turbine engine 100 to operate at highertemperatures NOx emissions can be reduced. The liner 12 made ofrefractory oxide ceramic material can replace current metal designs,which are cooled by impingement and film cooling and need hightemperature turbine alloys. The cooling of the ceramic liner 12 may bevia radiation to the metallic shell 13.

As discussed above, the bricks 15 are made of a refractory oxidematerial. Some examples of the oxide ceramic materials that can be usedto construct the bricks 15 are zirconium oxide, titanium oxide, aluminumoxide, mullite, combinations thereof, and the like. For example brick 15may be composition of SiO₂ and Al₂O₃. Preferably the brick 15 is aconglomerate of multiple phases, such as mullite and aluminium oxide.This type of oxide ceramic material is capable of being easily cast inorder to form the bricks 15 necessary for the formation of the gasturbine engine component.

FIGS. 3 and 4 show a close up view of bricks 15 that form the liner 12.The liner 12 is shown being used in conjunction with an interlayer 11that is located between the liner 12 and the shell 13. The interlayer 11may be a ceramic fiber mat 11. The ceramic fiber mat 11 is used in orderto provide the shell 13 additional protection from the heat as wellproviding a contact buffer zone between the liner 12 and the shell 12.The fiber mat 11 may also be used to dampen vibration. A further purposeof the fiber mat 11 is to minimize hot gas flow between the liner 12 andthe shell 13. Another purpose of the fiber mat 11 is to provide aninward spring force on the liner segments bricks 15 to ensure positivecontact under operating conditions.

The ceramic fiber mat 11 may be made of ceramic materials such asalumina, mullite, aluminosilicate, yttria alumina garnet, siliconcarbide, silicon nitride, silicon carbon nitride, molydisicilicide,zirconium oxide, titanium oxide, combinations thereof, and the like. Thefiber material used in the ceramic fiber mat 11 may comprise a non-oxidematerial. The fiber material may comprise ceramic fibers sold under thetrademark Nextel, such as Nextel 610, and 720 fibers. In addition, fibermaterial may be in any suitable form, such as a straight filament, abundle or a roving of multiple fibers, a braid, or a rope. The fibermaterial may comprise non-ceramic materials, including but not limitedto carbon, glass, polymeric, metal, or any other suitable fibermaterials.

Each of the bricks 15 has a first mating side 16 and a second matingside 18. The first mating side 16 and the second mating side 18 arecontoured so that they complement each other. First mating side 16 iscontoured outwards (i.e. convex) so that it is shaped to engage secondmating side 18, which is contoured inwards (i.e. concave). The pressuresexerted by adjacent bricks 15 when assembled cause the bricks 15 to matewith each other and remain in place. While the bricks 15 may mate witheach other in this manner, other means for engaging each of the bricks15 may be used, such as pins, grooves and other interlocking assemblies.

FIG. 5 shows a cross-sectional view of a pin 21 that can be used withthe liner 12 and shell 13 in order to securely attach the cone 10 to thegas turbine engine 100. The pin 21 may be secured at different receivinglocations 22 around the combustor basket 8. In the embodiment shown inFIG. 5 there may be three receiving locations 22 that are located 120°apart around the circumference of the gas turbine engine 100. Thelocations of the receiving locations 22 provide distribution for thecircular shaped cross section of the cone 10. The receiving location 22may be threaded so to permit the pin 21 to be threaded into thecombustor basket 8.

Other attachment means can be used to connect the pins 21, such asbolts, etc. Preferably the pins 21 are removable so as to permit easyrepair and replacement of the bricks 15. The pins 21 should be able toaccommodate thermal growth of the cone 10 while limiting movement of thecone 10 in the axial and circumferential directions. Additionally thepins 21 should be able to accommodate the weight that the cone 10 mayhave due to the use of the bricks 15, which may be heavier than othermaterials typically used.

FIG. 6 is cross sectional view of a cone 10 that uses a metal springs 23between the shell 13 and the liner 12. The metal springs 23 can be usedin place of the air gap 9 or the interlayer 11. However, in someembodiments combinations of the air gap 9, the interlayer 11 and themetal springs 23 may be used in the area between the liner 12 and theshell 13. The springs 23 are able to accommodate thermal fluctuations ofeach of the bricks 15 as the impacted by the heat of the working gases.In addition to being able accommodate movement of the entire liner 12 asit thermally fluctuates, the use of the springs 23 can also isolatemovement of one brick 15 versus movement of another brick 15 within theliner 12. The springs 23 may also be used to dampen vibration loads.Another purpose of the springs 23 to provide an inward spring force onthe liner segments bricks 15 to ensure positive contact under operatingconditions. The use of a metal spring 23 may also require cooling air tobe passed between the liner and the shell. As an alternative springs 23may be wave springs made of CMC. Having the springs made of CMC canenable the springs 23 to handle higher temperatures without the need forcooling air. The springs 23 may also be used to dampen vibration.Another purpose of the springs 23 to provide an inward spring force onthe liner segments bricks 15 to ensure positive contact under operatingconditions.

While embodiments of the present disclosure have been disclosed inexemplary forms, it will be apparent to those skilled in the art thatmany modifications, additions, and deletions can be made therein withoutdeparting from the spirit and scope of the invention and itsequivalents, as set forth in the following claims.

What is claimed is:
 1. A gas turbine engine comprising: a combustorbasket; a cone connected to the combustor basket, the cone comprising; aliner; and a shell surrounding the liner, wherein the liner is formedfrom a refractory oxide ceramic material and the shell is formed frommetal.
 2. The gas turbine engine of claim 1, wherein the liner has anouter surface and an inner surface, wherein the distance between theouter surface and the inner surface is greater than 20 mm.
 3. The gasturbine engine of claim 1, wherein the liner is comprised of a pluralityof bricks.
 4. The gas turbine engine of claim 3, wherein each of theplurality of bricks has a first mating side and a second mating side,wherein the first mating side of the plurality of bricks mates with thesecond mating side of another of the plurality of bricks.
 5. The gasturbine engine of claim 1, wherein the refractory oxide ceramic materialcomprises SiO₂ and Al₂O₃.
 6. The gas turbine engine of claim 1, whereinthe refractory oxide ceramic material is a conglomerate of mullite andaluminum oxide.
 7. The gas turbine engine of claim 1, further comprisingan interlayer located between the shell and the liner.
 8. The gasturbine engine of claim 7, wherein the interlayer is made of a ceramicfiber mat.
 9. The gas turbine engine of claim 1, further comprising ametallic spring located between the shell and the liner.
 10. The gasturbine engine of claim 1, wherein the cone is connected to thecombustor basket with a plurality of pins.
 11. A component for a gasturbine engine comprising: a liner; and a shell surrounding the liner,wherein the liner is formed from a refractory oxide ceramic material andthe shell is formed from metal.
 12. The component of claim 11, whereinthe liner has an outer surface and an inner surface, wherein thedistance between the outer surface and the inner surface is greater than20 mm.
 13. The component of claim 11, wherein the liner is comprised ofa plurality of bricks.
 14. The component of claim 13, wherein each ofthe plurality of bricks has a first mating side and a second matingside, wherein the first mating side of the plurality of bricks mateswith the second mating side of another of the plurality of bricks. 15.The component of claim 11, wherein the refractory oxide ceramic materialcomprises SiO₂ and Al₂O₃.
 16. The component of claim 15, wherein therefractory oxide material is a conglomerate of mullite and aluminumoxide.
 17. The component of claim 11, further comprising an interlayerlocated between the shell and the liner.
 18. The component of claim 17,wherein the interlayer is made of a ceramic fiber mat.
 19. The componentof claim 11, further comprising a metallic spring located between theshell and the liner.
 20. The component of claim 11, wherein thecomponent is adapted to be connected to a cone combustor basket with aplurality of pins.